A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection is between the fan case and the core engine includes three triangular-frame connecting members rigidly connected to the fan case at a fan case connection point, and to the core engine at a core engine connection point. The triangular-frame connecting members each are defined by two rigid legs which extend between the fan case and to the core engine, along directions each have a component extending radially inwardly and a component in opposed circumferential directions to each other. A plurality of non-structural fan exit guide vanes and the non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise. The non-structural fan exit guide vanes are rigidly mounted to at least one of the fan case and the core engine.

A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection between the fan case and the core engine includes a plurality of H-frame connecting members rigidly connected to the fan case, and to the core engine. The H-frame connecting members each are defined by two rigid legs which extend between the fan case and to the core engine, along directions which are generally parallel to each other. A plurality of non-structural fan exit guide vanes and the non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise. The non-structural fan exit guide vanes are rigidly mounted to at least one of the fan case and the core engine.

A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection is between the fan case and the core engine including a plurality of radial struts rigidly connected to the fan case, and to the core engine. A plurality of non-structural fan exit guide vanes and the non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise. The non-structural fan exit guide vanes are rigidly mounted to at least one of the fan case and the core engine.

A turbine assembly (1) comprising: —a plurality of turbine ring sectors (20) made of ceramic-matrix composite material, —a ring support structure (3), comprising an annular shroud (6), and in addition −a plurality of angular spacer sectors (70) together forming an annular spacer (7), said annular spacer (7) being, on the one hand, fixed to the turbine ring (2) and, on the other hand, fixed to said annular shroud (6), characterized in that said turbine assembly (1) comprises at least one air diffuser (8), which is configured to diffuse cooling air onto the radially outer face (212) of at least one of said turbine ring sectors (20), and in that said at least one air diffuser (8) is mounted by being nested on one of said angular spacer sectors (70), in a nested position.

A gas turbine engine including: a first turbine rotor assembly including a plurality of first turbine rotor blades extended within a gas flowpath; and a casing surrounding the first turbine rotor assembly, wherein the casing comprises an outer casing wall extended around the first turbine rotor assembly; a plurality of vanes extended from the outer casing wall and within the gas flowpath at a location aft of the first turbine rotor assembly; and a thermal control ring positioned outward along a radial direction from the outer casing wall, and wherein the thermal control ring comprises a body and a plurality of pins, and wherein the plurality of pins extend between the outer casing wall and the body.

An assembly includes first and second core gaspath walls. Each of the core gaspath walls defines a core gas path side and a non-core gas path side. The first and second core gaspath walls are arranged next to each other and define a gap therebetween. There is a seal arranged on the non-core gas path side that bridges over the gap to seal the gap. A spring device has a plurality of spring elements. The spring elements bias the seal against the non-core gas path sides of the first and second core gaspath walls.

A turbine stator includes a partition plate including an inner ring that extends along a circumferential direction, an outer ring that is disposed on an outer side of the inner ring in a radial direction, and extends in the circumferential direction, a plurality of nozzles that are disposed in the circumferential direction between the inner ring and the outer ring, and are configured to guide a fluid from an upstream side toward a downstream side in an axial direction, and an annular protruding portion, protrudes from the outer ring to the downstream side in the axial direction, and extends along the outer ring in the circumferential direction, and a casing surrounding the partition plate from the outer side in the radial direction, and having a contact support surface that is in contact with the annular protruding portion from the downstream side in the axial direction.

An assembly includes a gas turbine engine component and a plate. The plate is spaced from a surface of the component and generally conforms to the shape of the surface. The plate and component form a passageway that allows for passage of a secondary gas flow between the component and the plate.

An axial compressor comprises a plurality of compressor stages positioned axially adjacent one another within a casing. Each of the plurality of compressor stages comprise a rotor segment and a banded stator segment positioned axially adjacent the rotor segment. An annulus is formed between the casing and an outer flowpath ring of the stator segment. A vane of each stator segment comprises a member extending into the annulus.

An axial compressor comprises a plurality of compressor stages positioned axially adjacent within a casing. Each of the plurality of compressor stages comprise a rotor segment and a banded stator segment positioned axially adjacent the rotor segment. An annulus is formed between the casing and an outer flowpath ring of the stator segment. A pathway is provided that establishes an air flowpath between an entry pathway in a first stage and an exit pathway in a second stage.